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Does the flow over a wing decelerate upward, like air decelerates sideways with a swept wing?
In this answer, it states that the deceleration of air in the region of pressure rise over a wing acts in a direction perpendicular to the lines of relative chord, explaining how spanwise flow is formed.

To be precise: The acceleration and deceleration acts in a direction perpendicular to lines of the local relative chord.

My main question is does the air decelerate in the vertical axis too? Imagine taking the spanwise flow and flipping it 90 degrees so it was pointing up. In other words, I'm asking if there is a flow component facing upward in the region of pressure rise. (because the acceleration and deceleration of air acts in a direction perpendicular to the lines of relative chord, and the vertical axis would be perpendicular) Below is an image to convey what I mean about vertical flow component. (Feel free to ask for clarification in the comments, it's probably hard to understand what I'm asking.)
Here is an edit that I just did, because I now have another question : Does the air decelerating vertically contribute to stalls? (the air decelerating vertically towards the trailing edge)

Actually fascinating that, what is usually referred to as "downward acceleration", is being referred to here as "upwards deceleration", but OK!  Studying stalling onset of swept wings vs straight wings may help.  Interesting that the airflow is also "bent" by the wingtip vorticies.

@RobertDiGiovanni well I had assumed that if air decelerates sideways for spanwise flow, it would also decelerate upward.

In this photo, the streamlines just above the airfoil near the trailing edge are spaced more widely.  So, as long as there's no vacuum there (as there might be for hypersonic flow), the airflow there must be "moving up" relative to ambient, as the red arrow suggests.  That air has to come from somewhere.
A chord length or two above the airfoil, the spacing returns to ambient, so this upward flow has vanished by then.  Thus, the upward flow must eventually decelerate to zero.  Similarly, because the upward flow has to start from somewhere, it must begin with an upward acceleration.

Ah I see, thanks. I'm going to copy/paste question from comments on other answer to here just to get your point of view : Say you had a swept wing, that was experiencing spanwise flow. Would the vertical component of the flow that you described be less, because some of the energy needed to do that is being used for decelerating air sideways for spanwise flow? (if that doesn't make sense I'll happily clarify it)

So, given two wings, identical except that one is rectangular, the other swept, are you asking if the upward flow near the trailing edge is less for the swept one?  (Never mind the reason; just measure it in a wind tunnel, or pummel it with Navier-Stokes simulation.)

yes that is correct

From the picture it can clearly be seen that air does accelerate vertically. If not, it would have just continued to travel on a perfectly horizontal line. Instead, due to the presence of the airfoil, it get pushed upward and downward i.e. it gets accelerated upward and downward.

ah okay thanks! One small question that might be hard to understand : Say you had a swept wing, that was experiencing spanwise flow. Would the vertical component of the flow that you described be less, because some of the energy needed to do that is being used for decelerating air sideways for spanwise flow? (again, if that doesn't make sense I'll happily clarify it)

this is a late reply, but another little question popped into my head. Does the air decelerating vertically contribute to stalls?

Air decelerating along the rear portion of the airfoil is the one mainly contributing to the detachment of the boundary layer i.e. to the stall.

I see, thanks. I was thinking it might be the whole reason wings stall in the first place.


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Do slats and leading edge devices increase lift at a constant AoA and airspeed?
I've seen mixed sources on this subject. Some say leading edge devices allow the wing to obtain a higher AoA, and others say it flat out increases lift. For me, the latter makes most sense, because you're forcing more air to the upper surface of the wing, making lower pressure, therefore increasing lift?
Examples of different sources saying different things : "When deployed, slats allow the wings to operate at a higher angle of attack before stalling. With slats deployed an aircraft can fly at slower speeds, allowing it to take off and land in shorter distances." -Wikipedia
"The leading edge slats play an essential role in landing and in takeoff which tend to increase coefficient of lift and the stall angle. They are especially useful in takeoff which increase the lift production at a low drag penalty" -ScienceDirect.
((I believe that slats increase stall AoA, but this whole question was mainly about if you compared the lift of 2 planes, one with slats deployed, and the other with them retracted. All at the same AoA and airspeed.))
(Also, these sources kind of say the same thing in a way, but the Wikipedia one doesn't talk about them increasing lift at the same AoA and airspeed (for what I saw). I'm not sure if the ScienceDirect one does either, but that's what I interpreted it as.)

Leading edge slats mainly modify the pressure distribution over the forward and upper part of the airfoil which is, at least at high AoA, the most stall-sensitive part of the airfoil. In particular, the leading edge slats forces the airflow to remain attached until higher AoA in respect to the same wing with no slats extended. So the AoA can be increased further without stall therefore obtaining higher lift as consequence.
The following figure (source) represents in a good way the net effect of a slats deployment:

Note that, being the slats (at least locally) at basically a negative AoA, it is locally creating a negative lift and therefore it's global contribution to the lift is also negative: this can be also seen in the previous plot where $C_l$ is slightly smaller at the same AoA. Anyway the delay of the stall region more than compensate for this local phenomenon.

@Pilothead: I personally interpret "flat out increase of lift" as "extend the linear portion of liff"... but I agree that the wording in the original question isn't that unambiguous... I'm going to delete that reference and give a more general explanation.

Please keep the original graphs are well.  The effect of the slat depends a lot on wing design and Reynolds number.  Full scale aircraft generate lift most efficiently with the top of the wing (subsonicly).  Lower Re models must bounce it off the bottom to generate lift (but they can be made "scale" because of their lower weight to surface area ratio).  Since this site deals with full scale (mostly) those graphs are good.

@sophit this new graph is better l. Also since we don't actually disagree, perhaps you should remove the downvote.

I see, that makes sense, thanks. I'm going to copy and paste a question I had in the comments of a different answer to maybe get your view on it : One question about leading edge flaps : because they’re at an angle relative to the chord line, wouldn’t they force air upward, promoting separation? Basically the air is coming to the main airfoil at an angle (because of the LE flaps giving it an upward velocity vector) so it has to work harder to follow the contour of the wing.

Leading edge slats and droops decrease the lift at low AoA due to the downturned nose generating negative lift.
So if your "constant AoA" is lower than the angle of flow separation onset, lift will decrease.

Cambered wings generate lift at even negative AoA.  Slats increase the camber of the wing and lower local AoA.  But aircraft AoA "weathervanes" can be located on another part of aircraft, such as the fuselage.

This is true according to pg 420 of Sophit's Anderson reference at https://archive.org/details/FundamentalsOfAerodynamics5thEdition/page/n419/mode/1up

One can also lower the flaps a little too (when slats are deployed).  This increases camber even more, while maintaining (chordwise) AoA.

Adding the constraint "...All at the same AoA and airspeed ... " defeats the value of any comparison and makes the statement meaningless. The point of adding slats is to allow the aircraft to generate enough lift to remain in flight at a slower airspeed, in order to reduce takeoff and landing distance.
And the amount that AOA increases as a result of deploying slats or flaps depends on how you measure the AOA. Adding slats and flaps changes the chord line of the wing, and as AOA is arbitrarily defined as the angle between the chord line and the flight path, changing the chord line makes the change in AOA totally dependent on how the chord line has changed, and therefore meaningless by itself as any indicator of performance changes. If you deploy slats across the entire leading edge of the wing, and do not deploy flaps (the F-4 Phantom could do this), then because the chord line rotates leading edge down, the AOA decreases as a result. The maximum lift that the wing could generate, nevertheless increased, and the minimum airspeed required to generate any specific amount of Lift decreases.
This is because AOA is measured from an arbitrarily defined zero point (by the chord line) and comparisons in AOA are only meaningful when comparing measured AOA that are relative to the same base point. It would be like concluding that the temperature in Europe is always hotter than in America (based on the reported temperatures in the media there), without considering that they measure in Centigrade not Fahrenheit.

I think I see where you’re coming from, thanks. One question about leading edge flaps : because they’re at an angle relative to the chord line, wouldn’t they force air upward, promoting separation? Basically the air is coming to the main airfoil at an angle (because of the LE flaps) so it has to work harder to follow the contour of the wing.

Leading edge slats are themselves shaped airfoils, with a leading edge that points downward, and grabs air from below the leading edge of the rest of the wing, and lifts it and turns it to point backwards along the top of the wing, flowing parallel to the top of the wing. It re-energizes the airflow across the top of the wing. with the air that flowed under the slat between the slat and the leading edge of the rest of the wing.

Consider the effect of deploying the wing lift mechanism on the F-8 Crusader. Would you say that the AOA increases, because you only see the direct absolute effect of raising the leading edge of the wing up from it's current position, as it would if you did it while parked on the ramp? No, because what really happens is that the fuselage is lowered. In the air, the wing itself is the reference point from which meaningful comparisons should be referenced.

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Wind tunnel flow over surface vs. non-wind tunnel (atmosphere) flow question
If you take an infinitely thin sheet of metal and stick it out of the window of an airplane at exactly 0 AOA, will the surfaces of it become low pressure if you speed up? In a wind tunnel, there is the fan making high pressure (which converts to lower pressure because of the faster moving air as you go down the tunnel, I think). In real flight, there is no fan or anything to make pressure differences around the plane that I'm aware of, like there is in a wind tunnel. Why are flows in a wind tunnel basically the same as real flight in that case?
Another way to ask this. From the perspective of the ground to a plane flying, no air is moving that fast so Bernoulli wouldn't work, right? If you could see air molecules, they wouldn't be moving that much compared to you on the ground. (besides downwash and other obvious effects)
This also might help : If you're going faster density decreases. From the perspective of the ground to the airplane, the same amount of air molecules are colliding and interacting with the plane, as if you were moving slower. Why would density decrease around a plane in that case, at high speed? (ignoring the fact that as you go faster air molecules will bunch up and make higher density air)
This question might seem all over the place, but it was really hard to explain.
Last thing that I just thought of : Bernoulli works by ‘stretching’ out air by using low pressure (making the air move faster in the process). In a wind tunnel, the fan does that. In real flight, there is no such function?

Bernoulli's theorem only compares the energy of the same flow at different speeds and elevations. There is no "absolute" Bernoulli value, it is all relative. Like with energy itself.

Your questions follow a common theme.  You tend to use your intuition to try to work through the questions -- but without the required math and rigor.  Intuition can give guidance, but it can also easily mislead.  You tend to over-complicate things when you're trying to simplify them.
In this case, comparing a 'real' flow perspective to a wind tunnel perspective, is simply a matter of a reference frame transformation.  You should look into Eulerian vs. Lagrangian frames.  The equivalence of these perspectives is developed for all Newtonian physics -- whether it is fluid mechanics, or a car driving down the road.
In particular, this idea works for wind tunnels of all speeds -- from incompressible (nearly Mach zero) to hypersonic.  This tells us that variation of density does not matter (as the density is constant in an incompressible flow).  So you don't need to consider any ideas about density change when thinking about changes in frames of reference.
Furthermore, almost all flows of interest for aircraft can be considered continuum -- i.e. we do not treat the flow as a bunch of particles bouncing off of one another.  Instead, we treat the flow as a continuous medium.
The exception occurs in the very upper reaches of the atmosphere as you reach the edge of space.  There, you transition from a continuum flow to free molecular flow (where everything is based on particle collisions).  This idea is expressed by the Knudsen number $Kn=\lambda/L$.  It is the ratio of the mean free path $\lambda$ (the average distance a particle travels before colliding with something) to $L$ the characteristic length scale of the object in the flow.  When $Kn<0.01$, you have a continuum flow.  The mean free path at sea level conditions is less than 70 nm=$70x10^{-9}$m.  A 1mm gnat is about 150,000 times too big to be worried about non-continuum effects.
Engineers use particle collision theory to derive the perfect gas equation of state $P=\rho\,R\,T$ (and a few other things), but then we leave it behind and just take the equation of state with us.
Stop thinking about particle collisions and density changes (in incompressible flows).  They aren't significant to the answers you seek and are just getting you wrapped around the axle.

If you take an infinitely thin sheet of metal and stick it out of the window of an airplane at exactly 0 AOA, will the surfaces of it become low pressure if you speed up?

No.
In that condition the speed is exactly the same on the whole surface of the thin plate (apart in the boundary layer that we can disregard) so there's no change of pressure either. As soon as the AoA changes then the airflow on the surface of the plate experiences local changes in speed and in pressure, changes which are quite well described by the Bernoulli's equation.

In real flight, there is no fan or anything to make pressure differences around the plane

As just said, if the plate is at 0° AoA then the pressure difference around the plate is always 0, full stop. It doesn't matter if the plate at 0° is moving in low pressure air, high pressure air, in a wind tunnel, in water, in lava, in the airflow of an air dryer or in the martian atmosphere: pressure and speed around a flat plate at 0° AoA is constant, it doesn't change, and the relevant aerodynamic force is zero as well.

Bernoulli works by ‘stretching’ out air by using low pressure

Bernoulli is just an equation that, within its limitations, describes how reality works. In particular, it just mathematically translates the fact that if a body immersed in a fluid makes the fluid change its speed locally around the body, then also pressure changes locally, inversely proportional to the square of the speed's change.

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Why do boundary layers become more turbulent as they flow over a surface?
Why do boundary layers become more turbulent as they flow over a surface? This question originated from this one.
What makes the boundary become more turbulent as it flows over a surface/wing? My suspicion is that it's from the varying speeds over the boundary layer (which becomes greater over a bigger surface), but why would that cause turbulence?

As a flow progresses, tiny perturbations happen.  In some flows, these perturbations are quickly damped out as if nothing happens.  This is a laminar flow.  In other flows, these perturbations grow and cause more perturbations.  This is a turbulent flow.
Whether turbulence happens is therefore a stability problem.  Will the tiny perturbation stay the same?  Will it be damped out?  Will it grow out of control?
There are two reasons that turbulence increases along the length of a flow.  1) you've had more opportunities for random perturbations, so turbulence is more likely.  2) the character of the flow changes with the length of a boundary layer such that a formerly stable flow becomes an unstable flow.
The Reynolds number is the ratio of the inertial force to the viscous force on a fluid element.  I.e. the ratio of the effects of the fluid element's momentum to the effects of viscosity in and on the fluid element.  You can probably imagine that if the inertia greatly outweighs the viscosity that you will have a different characteristic of flow than if the viscosity greatly outweighs the inertia.  The Reynolds number increases along the length of a flow -- so the fundamental character of the flow (inertial vs. viscous effects) changes along the length.  The stability of the flow (whether perturbations grow or shrink) also changes along with it.

Oh I see, thanks. One question: Why does the character of the flow change along the BL? What makes it change?

@Wyatt Please re-read my answer.  The Reynolds number represents the ratio of inertial force to viscous force.  That ratio changes along the BL (because it is a function of the flow length along the BL).  At some Re, inertia dominates -- at a different Re, viscous dominates.  The flow fundamentally has different character.  It changes from stable to unstable.  Think about driving on an icy road.  For some speed and mass, you are stable -- but too much inertia and you'll go unstable.  What if the road is only wet?  What if it is dry?  The ratio of the inertia to viscous is what matters.

That makes a lot more sense. Up until now I didn't really understand how the Reynolds number worked, but you helped me understand it. Seems like a very useful tool for determining if flows will be turbulent or laminar, (or in between). Thanks

Also one last little question : is it possible for the boundary layer to not become more turbulent as it goes over the wing? In a more perfect situation though.

In an external flow, the probability of a flow to transition to turbulent will increase with flow length.  If the Reynolds number is sufficiently low, transition is impossible and the flow will remain laminar.  There are some boundary layer control techniques (such as boundary layer suction) that can substantially delay the transition to turbulence.  Internal flows (inside pipes, between concentric cylinders, etc) can have very different character and a different world.  Note that turbulence is an area that physics, math, and engineering does not yet have a complete theoretical model of.

What makes the boundary become more turbulent as it flows over a surface/wing?

There are so many people who know much more about this than I.  But let's just discuss this simply regarding low mach number wings and flight, say Reynolds numbers more than a million but less than 10 million.  Typically, two things contribute to turbulence:  a) intrinsic factors, and b) extrinsic factors.
Let's take a look at the intrinsic factors, first.  As the flow progresses over the wing from leading edge to trailing edge, the flow initially accelerates to a maximum velocity (about mid-chord, say somewhere between 30 and 60 percent of the chord length), and then decelerates to the trailing edge.  We call this initial acceleration a favorable pressure gradient, and the deceleration region an adverse pressure gradient.  The favorable pressure gradient promotes the growth of the laminar boundary layer.  At the initiation of the deceleration phase, however, the laminar boundary layer is destabilized and forms a laminar separation bubble, which itself increases drag. The laminar boundary layer is destabilized by the loss of the favorable pressure gradient and becomes turbulent as the flow reattaches and progresses against the adverse pressure gradient. No longer is the velocity of the boundary layer zero at the surface of the airfoil, as it was with the laminar boundary layer, but the energized flow is now actively moving against the surface.  As this turbulent boundary layer progresses against the adverse-pressure gradient, the turbulent layer becomes thicker and will increase skin-friction drag with the airfoil.  The adverse gradient can cause the turbulent layer to separate resulting in a marked increase in pressure drag. Momentum from the aircraft is now transferred into the turbulent airstream separating from the wing. The wing has stalled.  However, this is not always true.  Airfoils can be designed in such a way that the shape of the airfoil controls the gradient of deceleration against which the turbulent boundary layer must progress.  In fact, the shape of the airfoil profile through this region can be designed in such a way that the boundary layer becomes infinitely separating and produces no skin friction or pressure drag as the flow separates.
What are extrinsic factors?  Bugs. Dust, surface imperfections, dents, rivets, rain, the propeller.  Any of these items can trip the laminar boundary layer causing it to become turbulent in the favorable pressure gradient of the airfoil. This increases drag.  As this flow accelerates, the turbulence generated thereby becomes fully energized and will thicken precipitously as the flow progresses through the adverse pressure gradient.  Nevertheless, this process can be propitiously used to advantage.  If a trip strip is placed on the surface of the airfoil in the favorable pressure gradient region just ahead of the adverse-pressure gradient where the laminar boundary layer will separate, the tripped laminar layer will become turbulent and remain attached.  Overall, tripping the laminar layer in this way will reduce separation-bubble drag and promote the onset of turbulence which can promote maintaining attachment of the turbulent boundary layer.
But what about the entire airplane?  In today's world, every attempt is made to keep surfaces smooth, contours smooth, turbulence smooth.  The separation process can happen anywhere along the aircraft itself, and promote increased drag.  Careful shaping of the aircraft can resolve this issue, but the design and construction work takes time, patience, and care.  Take a look at this link and watch this series about what goes into making smooth work to one's favor.


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What happens when 2 air streams, one going faster but the same direction, are beside each other?
Inspired by this video, what happens when 2 air streams beside each other going the same direction, but one going faster, interact? Do they make turbulence and or vortices as said in the video?
I’m sure turbulence is created because of what the video said, but why is it created, and is it possible for vortices to also form?
(When I say beside each other, I mean the air streams right beside each other, not separated by any gap)

"going the same direction, but one going faster": For this matter, the direction doesn't count, only the absolute value of the velocity difference.

Look at this image of a turbine core exhaust as it mixes with the slower airflow that bypasses the core. You can see the shear layer from the high speed difference grow into small scale instabilities, then become larger as mixing progresses. There are vortices at numerous locations in the flow.
Mixing occurs as the fluids interact with each other. Right at the nozzle the flows are already mixing because molecules even in "stationary" fluids are moving in all directions and easily cross the boundary between the flows, although in a very thin layer at this point. As molecules from the slow flow collide with those from the fast flow and vice versa, the thickness of the layer of colliding molecules continues to grow until the energy differential has been eliminated. In a fully mixed flow all the molecules are back to being in a "stationary" state; there are no longer two discernable flows.

Outstanding picture.  On can see that once the faster flow loses velocity (and the surrounding air gains velocity), some serious mixing (turbulence) occurs.  Near the outlet, the high velocity air will draw surrounding air into the stream, but eddys will not form until further downstream.  "There are no longer 2 discernable flows", but rotational energy in the form of vorticies can remain.

Two streams next to one another form a boundary layer -- just like a boundary layer between a flow and a solid wall.
In the 'middle' of a flow, nearby velocities must be parallel.  Hopefully this is intuitive, if they aren't parallel, then the conservation of mass is violated.
When the velocities are parallel, but different magnitude, a shear force exists.  Imagine a block (cube) of flow.  One face has velocity $V$, the opposite face has $V+\Delta V$.  If you think of the cube as a block of jello, it will shear due to the $\Delta V$.
Those shear forces related to $\Delta V$ are viscous in origin.  It is the friction between the 'layers' of flow that cause the shear force.  The shear force tries to slow down the fast stream -- and speed up the slow one.
In the middle of most flows, these $\Delta V$ are all small and we can generally think of the bulk of the flow as inviscid.  These shear forces are generally small everywhere.
At a wall, the $\Delta V$ is large -- the velocity at the wall is zero, so a strong boundary layer forms where the effects of friction are strong and turbulence may appear.
When you have two streams of flow that meet (and are parallel), a shear layer will form.
Previously, I have ignored viscosity and I have described this to you as a slip line.  I.e. if we let the flow 'slip' over one another, we allow the $\Delta V$ to exist without viscoscity.
When we consider that friction exists, we observe the formation of a shear layer.  Like a boundary layer on a body, it will grow in thickness as it moves downstream.  It is trying to speed up the slow flow -- and slow down the fast flow.

Ah I see, thanks for your answer. So a shear layer is created like you said, but does that necessarily mean that turbulence is created?

No, you can have a laminar shear layer under certain circumstances.

Oh okay yeah that makes sense. Also, I’m assuming that the Re number has something to do with it?


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How do thicker airfoils give better stall characteristics?
Why do thicker airfoils have better stall characteristics? They make more lift, but I don’t see any reason for them to have better stall characteristics. Also, I heard a rounded airfoil nose will help with this, why is that?
Edit : better stall characteristics meaning that it has a progressive loss of lift and doesn’t loose lift all of the sudden. Also the pitching moment changes smoother.

Please define what you mean with "better stall characteristics"

@sophit oh I actually forgot to add that part, thanks for reminding me!  Fixed.

@sophit Also do you know why this question is getting downvoted? I don’t see a problem with it.

No idea why the downvote, it would be nice to get first a suggestion and then possibly a downvote ♂ In my opinion is a legitimate and interesting aerodynamics question

@sophit exactly, I wish people would comment when they downvote. I feel like not commenting almost defeats the purpose of downvoting, as you don’t know what to improve.

"I don't see any reason for them to have better stall characteristics"  Time to do some reading about (subsonic) airflow around a sharp and rounded leading edge surface.  Start with a flat plate and see what one would do with the shape of the leading edge to keep airflow attached as AoA increases.

@RobertDiGiovanni That makes sense, thanks. That almost seems like it could be its own answer!

For airfoils used in GA, "thicker" and "rounder nose" are identical. The thicker it is, the larger the radius of curvature at the nose must be.
A rounder LE will

increase the range of angles of attack at which the boundary layer stays attached, and allow for a more gradual transition between normal flow and stall

by reducing the pressure gradient that tries to detach the boundary layer from the airfoil.
-- https://aviation.stackexchange.com/a/58452/31425


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Why do shockwaves extend past the body that created them?
Why do shockwaves extend past the body that created them? As seen in this photo, the shock doesn’t stop in the air the plane is effecting, but continues on. I always assumed it was high pressure air from the shock extending out, but now I’m not too sure.

Pretty pictures

This is more of a follow up question of this question. I really just want to confirm what I think is correct. Thanks!

You aren't going to believe this answer, but it is true!
Think about a simple 2D diamond airfoil.  We will think of this as an inviscid problem for now.

The flow in region 1 is all parallel.  Start by considering the streamline just above the centerline.  When that streamline encounters the airfoil, it must turn by angle $\epsilon$.  This abrupt turning is what causes the shock.
Now consider the 'next' streamline.  In region (1), it must be parallel to the first streamline -- parallel to the centerline.  This streamline never encounters the airfoil.  Instead, it encounters the inner streamline when that streamline has been forced to turn.  The inside streamline turns, so the next streamline must turn.
Consequently, all the flow in region (1) is parallel.  Also, all the flow in region (2) is also parallel.  The shock exists to abruptly turn the flow by the angle $\epsilon$ -- it also causes a pressure jump and other changes in properties.
If this was all there is to it, region (1), the shock, and region (2) would go on to infinity!
Every streamline needs to turn because the streamline 'inside' of it turned.  Every streamline acts like the rigid wall along the centerline and airfoil surface.
Next, lets consider some things that make this break down a bit.
Between regions (2) and (3) is an expansion fan.  The fan acts to turn the flow the other way (convex corner) and it accelerates the flow.  Although the streamline at the body will turn abruptly, the streamlines further out in the expansion fan will turn gradually.

Notice that the angle of the 'start' of the expansion fan is such that it will eventually intersect with the oblique shock.
This is where things get really interesting.
Notice that most analysis, drawings, and other depictions stop with the very nearby vicinity of the flow like the first image.
When the expansion fan meets the oblique shock, things get interesting.  Recall that if the shock turns the flow 'up', the fan turns the flow 'down'.  When the meet, it has the effect of allowing the 'next' streamline to turn a little less.  Instead of turning $\epsilon$, it turns just a tiny bit less, and the next streamline turns less, and less for each one out.  This creates a region with a curved shockwave.  Each shockwave further out is a little weaker because it doesn't have to turn quite as much as the one before it.
Another bit of reality is that airplanes are 3D, not 2D.  So, airplanes behave more like the cone shock equations not the wedge shock relations.
We can think of conical flow as being axis-symmetric.  A streamline in 2D is really a cylindrical stream surface in 3D axis-symmetric flow.  However, when a flow turns 'out' from the centerline, the area available to it increases.  This is the fundamental reason why cone shocks are different from wedge shocks.
We call this a 3D relieving effect.  Streamlines further from the body have more room to move to.  This means that a 3D shock is really very different from a 2D shock.

That actually makes a lot of sense! I never had thought about it that way. Thanks for your answer!

Also something I’ve been thinking about: Because shocks aren’t going to expand perpendicular, is it possible for them to extend parallel? They obviously do, but is it possible for them to because of them being high pressure?

I really don't understand what you mean by 'extend parallel'.  Shocks are not because of the pressure.  Shocks happen because of the turning -- the pressure jump is a result of the shock.

Extending parallel meaning outwards. Basically the same thing as the original question was asking. I understand that the high pressure is a result of the shock, but does that high pressure, in any circumstances, help the shock extend outwards? I'm mainly referring to the normal shock on the surface of a wing.

Usually the normal shock on the surface of a transonic wing only has very short extent.  This is because the region of supersonic flow is limited to a bubble on the upper surface of the wing.  Once you get to the edge of this supersonic bubble, there is no reason to terminate the bubble with a shock.  This is in contrast to the supersonic aircraft -- where the supersonic freestream is everywhere.

Ah I see, I think I get what you're saying. Thanks again!

Late comment on here, but I was thinking, what will stop the shockwave from extending to infinity? (Considering there’s no expansion fan) I tried looking this up and all I got was stuff about car shock absorbers

There is an expansion fan.  And in general, the shock does extend to infinity.  That is why the shock around a 50 foot long aircraft will hit the ground as a sonic boom when the aircraft is lying at 50,000 ft altitude.

Ah okay. So after the expansion fan, does the shock keep its strength until it reaches the ground?

Shock propagation to the ground is a complex thing.  The shock's strength will change as it goes through different layers of the atmosphere.  Temperature changes are most important, but humidity and wind also have an effect.  But basically yes, near the aircraft, there are different interactions that change the shock.  Once you get to some mid-distance away, the remaining shock structures are parallel and won't interact anymore (because they don't collide).  Those tend to extend until they reach the ground.  They also spread out to the sides.

I see. So basically weather, temperature, and wind changes will change the flow turning aspects changing strength?

(strength of shock)

Yes, not in a major way, but enough that if you work for NASA and you are doing detailed calculations about how a sonic boom from an overhead aircraft will be experienced on the ground -- you care and it matters.


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What happens to tip vortices in a sideslip?
Say you're in a right sideslip. What happens to tip vortices? I'm fairly sure that it would just reduce the strength of the vortex (for the left wing), because there is a force trying to keep the high pressure under the wing, if that makes sense. That force is the freestream because the wing is at an angle.
The right wing vortex would probably grow in strength, because that same force is helping the air escape from under the wing. (if I had to guess)

All of this with just a straight rectangular wing. No winglets or anything.

Think about what happens if the wing has dihedral, anhedral...

I have seen some info about this, about the existence of wingtip vortices on a flat (no-dihedral/anhedral) wing causing either an anhedral-like or dihedral-like effect (roll torque) in a sideslip.  Will try and locate and come back with an actual answer.  It is a good question.

@quietflyer oh okay great! Thanks

Desktop Aeronautics has a paper out that can be found on the internet.  There is a rolling motion cause by differences in wingtip airflow patterns when the wing is "oblique".  Apparently (good for slips), the lead tip tends to roll down, somewhat canceling dihedral effect.

Wake (not tip) vortices will stay as strong as they were in straight flight and be of equal strength on both sides. Since neither the aircraft nor the wake are rolling but stay level, lift must be balanced. Lift differences from dihedral effects will be canceled by aileron deflection as appropriate to maintain level flight. The wake vortices will move away from the airplane at the sideslip angle since they stay with the local air.

@PeterKampf -- I guess one way to look at the question is how does the lateral shift in  the tip vortices affect the aileron inputs required to balance the aircraft in roll.  Once those aileron inputs are being made then yes I suppose both vortices will be equal in strength or nearly so.  Your comment more nearly addresses the actual question which didn't really ask specifically about how the vortices affect the net roll torque created by slip (in the absence of aileron correction to maintain the bank angle)-- guess I just jumped right to the roll torque issue because I find it interesting.

The following quote is from "Oblique Flying Wings: An Introduction and White Paper" from Desktop Aeronautics, Inc., available at https://www.desktop.aero/library/ofwwhitepaper.pdf.

2.2.2 Effect of Oblique Sweep on Spanwise Lift Distribution
A well-known effect of wing sweep is the variation of induced downwash
along the span from the trailing wake that produces an additional lift
distribution characterized by increased loading on the aft wing and
reduced additional lift on the forward wing. For a wing with no twist
or bend, this results in a significant rolling moment, tending to roll
the forward wing downward. There is also a yawing moment from the
asymmetrical distribution of induced drag that tends to unsweep the
wing.

Keep in mind that by "wing sweep" the authors are talking about obliquely sweeping both wings together as a unit, like in the photo in this other answer to the present question.  We have essentially the same effect when we sideslip.  The direction of the roll torque generated is toward the "forward" wing, which in a sideslip is the "upwind wing".  In a steady-state non-turning sideslip this would also be the "lowered" wing.  This is an anhedral-like effect.
In the paper they recommend adding a small amount of dihedral to the outboard portions of an oblique-wing aircraft to counteract this, with the amount of dihedral being dependent on the lift coefficient that the aircraft is intended to fly at.
I'd suggest that this effect due to "the variation of induced downwash along the span" falls within the scope of the question's focus on "tip vortices".
The anhedral-like (upwind) roll torque generated by an unswept, zero-dihedral wing in a sideslip was also the subject of a post on an aerodynamics subforum of the"RCGroups" model-airplane on-line forum by a prolific, knowledgeable contributor with the username "ShoeDLG".  Some excerpts from his post:

I used a Vortex Lattice Method to explore the impact of various
Radian Pro wing design features on apparent dihedral effect.
The cases I looked at (all without any twist):

Untapered / Unswept / Zero Dihedral Wing (flat AR 10 rectangular wing)..."

and then

The results are shown in the first attachment. A couple of
observations:

A flat rectangular AR 10 wing has negative apparent dihedral effect at positive AOA and positive apparent dihedral effect at negative
AOA..."

and then

The dihedral effect associated with an untapered, unswept, zero
dihedral wing is interesting. The span loading for an AR 10
rectangular wing at 5 degrees AOA and 20 degrees sideslip (to
exaggerate the asymmetry) is shown in the second attachment."

These conclusions are consistent with the "Oblique Flying Wings" white paper-- during a sideslip, on an aircraft with flat (no dihedral or anhedral) wings with no sweep, the lift distribution is changed in a way that creates an "upwind" roll torque-- i.e. an anhedral-like effect.
Here are the two attachments:

However, a web page called www.flightlab.net by Bill Crawford came to the opposite conclusion, stating that during a sideslip, tip vortices shifted in a way that increased the angle-of-attack of the upwind wing and decreased the angle-of-attack of the downwind wing, contributing a dihedral-like roll torque component, toward the downwind wing.  A full page of text devoted to the topic, along with an illustration, may be seen on page 14 of his generally excellent PDF "Flightlab Ground School: 4. Lateral/Directional Stability", now viewable only on this link on the Wayback Machine.  This content is well worth reading through.
Crawford goes on to reach the following conclusion: (footnotes added)

"Since downwash strength is a function of CL, pulling or pushing on
the stick will affect roll moment due to sideslip in a manner similar
to the swept-wing example already described.1 (Our trainers’
rectangular planforms tend to promote strong tip vortices. Other
straight-wing planforms with different lift distributions might not be
as effective.)"

followed by:

From all the above, an under-appreciated yet nevertheless great truth
of airmanship emerges: For a swept or a straight wing, pulling the
stick back tends to increase rolling moments caused by sideslip (and
by yaw rate), pushing decreases them.

(In the specific case of an unswept wing, Crawford is obviously assuming that there is some aspect of the aircraft's overall geometry, such as a high wing position, or actual dihedral, or perhaps just the tip vortices themselves, that results in an overall dihedral-like roll torque during a sideslip.  The comment above wouldn't make sense for, say, a low-wing airplane with no sweep and a great deal of actual anhedral, because then the effect that he is claiming for the wingtip vortices, which are most pronounced at high Cl, would be acting in the wrong direction.)
So Crawford is reaching the opposite conclusion on the effect of tip vortices on the balance of roll torque during a sideslip than the other sources cited above-- he is claiming that they contribute a dihedral-like roll torque component, toward the downwind wing.  I can't account for this discrepancy at present.  Is it possible that Crawford's conclusions are valid for wings of a certain geometry, such as the low-aspect ratio wings with broad wingtips found on many trainers?  Or is he simply mistaken in the idea that tip vortices, and the associated larger upwash / downwash flow field around the wings, tend to contribute a dihedral-like, downwind roll torque component during a sideslip? Should his observation of increased rolling moments due to sideslip at high lift coefficients versus at low lift coefficients, even in unswept wings, be explained by other means entirely?
Footnotes:

I.e. adding to the dihedral-like roll torque that would be created by sweep

Obviously I've jumped right to the topic of how do the tip vortices and other associated aspects of the flow field tend to generate a roll torque during a sideslip.  That wasn't actually the focus of the question but hope it is somewhat helpful and interesting.

Excellent resources.  Crawford states that rolling away from the side slip (lead wing) theoretically would happen even if dihedral was not present.  Oblique wing testing data may have shown other wise.  Super +1 up to that point.  Data presented by "ShoeDLG" supports (extensive) oblique wing studies.

Thank you for your answer! One question : what is meant by “tip vortices shifted in a way that increased the angle-of-attack of the upwind wing and decreased the angle-of-attack of the downwind wing”? I guess I’m asking why Crawford thinks the upwind wing has an increase in AoA, and the downwind has a decrease.

That is really interesting! Your clarity has provided valued insight in answering this question.  Thank you!

@Wyatt -- if you follow the "wayback machine" link and look at his actual figure on page 14 of the PDF it might help; I was trying to puzzle it through as typing the answer and couldn't really clearly see why he was coming to that conclusion.

(ctd) Looking at it again now I do notice that he has drawn the curved arrows representing the tip vortices larger for the left (downwind) wing than for the right (upwind) wing so maybe he is suggesting that by being shifted more toward the wing, the upwind tip vortex gets diminished in intensity, i.e. the wing itself interferes with the vortex much as a winglet or endplate would? If that's what he was thinking I wish he would have spelled it out explicitly in the text as that point would bear directly on your actual question more than the resulting roll trim effct that I was focusng on.

(ctd) Anyway if that actually happened, then the resulting change in downwash velocities over the wing would seem to be expected to contribute to a dihedral-like downwind roll torque, like he is saying, but the results given in the White Paper on oblique wings and the results of "ShoeDLG's" vortex lattice analysis suggest that the overall complete effect of the shift in the overall flow field due to sideslip actually causes the opposite effect.

(ctd) Both could be true at once: it could be true that the energy contained in the immediate area of the tip vortex gets decreased in the upwind wing, yet the net overall effect of the overall shift in the flow field ends up contributing a net anhedral-like upwind roll torque nonetheless.

(ctd) I may end up adding some of the above conjecture as a footnote to the answer later.   Possibly along with a screenshot of the figure on page 14 of the Wayback Machine link.  Possibly along with a little note at the start pointing out that this answer is focussing on one specific aspect of how sideslip affects the tip vortices and other associated aspects of the flow field, namely their effect on roll trim.

@Wyatt -- btw I have noticed that another RC glider, the basic "Radian" which lacks ailerons and has much more dihedral than the "Radian Pro", in sustained inverted flight is much more roll-responsive to rudder inputs with the nose more earthward (wing at small negative lift coefficient) than with the nose more skyward (wing at a large negative lift coefficient).  The opposite of what you'd expect from a wing with no dihedral or anhedral, but lots of sweep.  In inverted flight, from an external viewpoint, the wing has anhedral not dihedral,

(ctd) and depends on that anhedral for the rudder-roll coupling to work, so a dihedral-like roll torque contribution from the tip vortex displacement like Crawford was claiming could be one possible explanation.  I can think of some other things too though, like high-speed flight causing some of the anhedral to be removed as the  rather flexible wing "unbows" due to the way that the washout moves the (negative) lift distribution further outboard when the (negative) a-o-a is lower compared to when the (negative) a-o-a is higher.

ctd)  For example at very slight neg a-o-a (very high speed inverted flight), one might expect only the (washed-out) tips to be making any (negative) lift at all, while the inboard sections are making zero lift or very slight positive lift, which would tend to "unbow" the wing.  Would want to use an onboard camera to rule out that this is happening, before claiming that the observed results actually support the effect Crawford is claiming.  Or replicate tests with a much less flexible wing -- and possibly with zero washout.

@PeterKampf -- comment above beginning "Looking at it again now I do notice that he has drawn the curved arrows representing the tip vortices larger for the left (downwind) wing than for the right (upwind) wing" bears on your comment (under question) suggesting that the vortices will stay the same in intensity.

@quietflyer thanks for the detailed response! If I had to guess, I’d say that Crawford might be right when the wing planform is a certain way, like said in your answer. Particularly when it’s a low aspect ratio wing.

In consideration of level flight, if we think about dihedral, the wing yawed into the wind would develop increased lift relative to the other wing, and we would expect the lift-induced vortex on the wing yawed into the wind to be stronger than on the other wing.  This would occur in gliding flight and can be easily demonstrated in general, even though this is not exactly level flight.  However, in powered flight with a single-engine propeller on a mid-wing aircraft, the propeller slipstream may increase lift on the down-wind wing more than dihedral would increase lift on the upwind wing in a side slip.  The consequence will likely cause a roll in the direction of the upwind wing. Now the question remains, in this circumstance which vortex would be increased?  If we consider the total-potential circulation to be more strongly shifted toward the down-wind wing as a consequence of the propeller slip stream, then we would expect that the down-wind wing would have the stronger tip vortex.  Nevertheless, this result is counterintuitive and seemingly unexpected.
There are relatively few in-depth resources discussing in detail this aspect of flight.  quiet flyer will likely have a more informed perspective.

Or do you mean something mostly like this? Robert T. Jones, at NASA, rather extensively investigated this aspect of flight.  Various summaries and perspectives of his work in this area are readily available on the internet, particularly via NASA.  This particular configuration of the wing is termed oblique. In this case, for level flight, there is no apparent asymmetry in the relative distribution of lift.

A crucial difference between slip and oblique data is that slip data (for roll behavior) must include(net) dihedral effect from  relative wind on wing, fuselage and vertical stabilizer, whereas oblique tests only change the orientation of the wing.
This allows us to isolate the effects of oblique sweep on roll behavior and analyze them more accurately.
It is possible the "upwind" or "reverse swept" wing's airstream flow into the fuselage degrades the low pressure area over the wing, but the "sweep" of the downwind wing's airstream flow helps preserve it, resulting in a roll (from lift imbalance) into the upwind wing.
This is supported by winglet studies and bird-like aircraft design.
Amazingly, the data supports 2 major features which could help make the design viable:

Rolling tendency can be controlled with +/- dihedral in design
Since rolling effect is increased by increase in AoA, it would
make sense to straighten the wing out at lower speeds (when AoA
must be greater) to mitigate the effect.

Other concerns are with fluttering tendencies of reverse swept wing tips, which, if overcome, could lead a path to a practical, economical way to increase to flight envelope of an aircraft into the transonic, and even supersonic regimes.

I don't see the term "tip vortices" or singular "vortex" in your answer.  So what happens to them?

@MichaelHall the relative wind may push the vortex further into the upwind wingtip and further away from the downwind wingtip.  This is supported by the windflow across the upwind wing into the fuselage.  Many streamlined designs have tapered, swept tips and/or winglets to lower vortex interference with the upper wing low pressure area.

Shouldn't that be included in your answer?  ;)

@MichaelHall that which interferes with lift is higher pressure from under the wing curling over the top of the wing.  The vortex proper forms from this and inflow of surrounding air into the downwash sheet.  Maybe semantics, but the lift killer near the upper wing surface may be a half (turn) vortex?  Notice approaching stall, higher pressure can "leak" in from the trailing edge too.  All this may not be in the neat form of vorticies.  Interestingly, with vortex generators (and delta wings), a stable vortex can actually help preserve lift.

You aren't getting the hint....

@MichaelHall please read section 2.2.2 (page 16) in the reference linked in quietflyer's answer.  There the cause of rolling tendency (mentioned in the question) is discussed.

From my comments above, what makes you think I might be confused about rolling tendencies?  Again, you aren't getting the hint... I'm telling you that "the relative wind may push the vortex further into the upwind wingtip and further away from the downwind wingtip" is the answer. It is direct, simple, complete, and lacks nothing. That doesn't mean you can't supplement it with further discussion, but as you often do, you talk in circles around issues related to the question without actually answering it.  Would you like me to edit your answer to add that sentence at the top?

@MichaelHall I don't think you are confused.  I do think that "vortex" and its effects are vastly overrated and misunderstood.  Vortex creation can actually help lift in some cases.  I choose for focus on effects unique to the oblique wing which (from references) touch on the body of the question (pressure differential-->rolling).  Sort of amusing that my answer sought to clarify quietflyer's, and yours sought to clarify mine.  This is positive.

What happens to tip vortices in a sideslip?

Going for brevity, and with thanks to another member here:
The relative wind may push the vortex further into the upwind wingtip and further away from the downwind wingtip.

Might that not tend to diminish the strength of the vortex on the upwind side?  Currently addressed to some extent in some comments under my answer, may move into actual answer later--

@quietflyer, perhaps it would, but I don't intend to supplement my answer with further discussion at this time.


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How are shockwaves able to refract?
How are shockwaves able to refract? As said in this,

When two shock waves collide, they interact with each other and produce complex patterns of compression, rarefaction, and reflection. The resulting behavior depends on the strength, speed, and direction of the shock waves, as well as the properties of the medium they are traveling through.

I understand how they might be able to pass through each other, but why would they refract? Light and optics are a bit different than aerodynamics at supersonic speeds I'd imagine.
Also, refracting meaning when a shockwave refracts off of another shockwave. I assume they have to be different densities to refract off of each other.

I'm not the down voter, but while the topic is interesting, as a Stack Exchange question this is currently low quality. If you hold a cursor over the down vote arrow the tooltip suggestion for that action is "This question does not show any research effort... Downvoting is exactly the correct action when a question does not demonstrate any effort to look for an answer first, before posting here.

Also, do you mean a shock wave refracts around higher/lower density air some distance away, or another shock wave, or light is refracted by a shock wave? It's not even clear what you are asking. What would the density profile of this shock wave look like, roughly? and How is the camera positioned for these plane-to-plane Schlieren images of shock waves? and What do the components of a shockwave's visual artefacts represent?

@uhoh I am mainly taking about the effect where shockwaves can refract off of each other.

@uhoh also how might this question be "low quality"? Besides missing the part explaining what I meant by refracting (added it just now), I fail to see how this is low quality.

Ah that's much better! +1 So refraction happens at gradients/interfaces, so probably what's most important is the angle between the two shock wave fronts, if it happens at all. For normal sound waves, the air response is linear so two waves pass through each other without affecting each other. But the shock wave response will be very non linear, and each can certainly interact (and maybe refract) the other. Interesting phenomenon! Add a link suggesting at least a little prior research (e.g. where you first read that shock waves can refract) and your question will be even better received.

Oh, from a Stack Exchange community point of view, short questions, questions without research and questions that have ambiguity (are unclear) are all considered low quality. It's the question post itself - how it's written, not the underlying question. You've already made a big improvement with that edit! Now add some kind of link referencing refraction and the perceived quality of the post will be even higher!

@uhoh I see. I was going to add something like a picture of it, but I personally couldn't find anything that looked acceptable. After a bit of research, I found where I initially got this idea from, adding it now!

Do you have any flat sandy beaches near you?  Have or a lake where you can observe waves and ripples interacting with one another when they intersect?

@MichaelHall do low amplitude gravity waves (i.e. ripples on a liquid surface under the influence of gravity) really interact with each other, or like normal sound waves do they pretty much pass through each other without interracting? (i.e. linear superposition -there are waves and ripples (and ducks!) in the first image and caption in that article. I don't think it's a good analogy for shock waves which are extremely high amplitude and very nonlinear.

@uhoh, perhaps there are significant differences that would make it a poor comparison.  Just asking in case observing waves interacting may shift perspective enough to create insight.

The quote you cite does not say refraction.  It says rarefaction.
Refraction (of light) is when light bends because it transitions from one material to another -- say air to water.
I do not think people talk about shocks refracting -- if you have an example, please post it and ask specific questions about it.
I found this image of a refracted shock.  Here, an explosion occurred at the 'Source' above the ground.  The initial shock spreads out spherically.  When  the shock hits the ground, there is both a reflection from the surface of the earth -- and also a transmission of the shock into the earth.

The shock will travel differently through earth than through air -- at a minimum, the speed of sound in a solid (or liquid) is much different than the speed of sound in a gas.
While this is a fascinating problem, I don't think it is what you're actually interested in.
Rarefaction is a process of becoming more rare -- spreading out -- weakening.  Shocks can interact in a way that they weaken one another.
Instead of using the words Rarefaction, I believe I usually use the term 'Expansion Fan'
Here is an image of two shocks interacting that cause an expansion fan -- I think this could also be called a rarefaction zone.

As some of the criticism alludes, this question shows that you've put in very little effort.  You're asking the answerer (us) to do more effort than you did.  That is generally bad form.  It is much better for you to come to us with small specific details after you've conducted extensive study.
Also, I saw that you posted this in Physics Stack Exchange.  I believe it is bad form to cross post identical posts at the same time.
Unfortunately, there is no 'Aerospace Engineering' stack exchange.  For the most part, Aviation SE is willing to tolerate us asking and answering Aerospace Engineering questions here, but we do occasionally get asked 'Is this Aviation'?
That said, Physics SE is not the right place either.

I see. So I understand that part, but if you look at the right diagram on this page from NASA, the shocks seem to bend as they cross. Is this considered rarefacting? - my usual way of asking questions is I will do my research (maybe not as much as I should), and then if I have specific questions I’ll come here and ask without providing evidence that I have done such research. I try to keep the question short and simple, but maybe I should

Start providing the research.

I rarely ask the same question twice, but sometimes I feel P.SE is more suited for some questions more related to physics (not that this one even is this time). Shouldn’t have done it, but at the time I thought it was a good idea.

One last thing: if shocks going into the ground refract, why wouldn’t they do that going through other shocks? The formula for sound speed is $c=\sqrt{\frac Kρ}$. If density is higher, (in shockwave) wouldn’t that change the speed of sound, making the shock refract?

(If the density was higher in the shock that a different one was going through)

Well actually never mind about that part; the shocks make destructive interference so they don’t have time to refract.

The figure at the right in your image from NASA.  Those shocks bending is not called rarefacting or refracting.  That is just bending a shock.  These two examples have great descriptions in the text on that NASA page.  That kind of thinking is exactly how you should approach problems like this.

I see. Is this problem related to this answer of yours in any way? It seems like the reason shocks extend outward would have something to do with why they bend when crossing. (Asking because if they're related it would definitely make sense)

YES.  Shocks are always a response to the boundary conditions.  The walls force the flow to turn, the shocks are how that happens.  The cool thing about regions 3 and 4 in the case on the right is that the shocks in that case find the solution that causes regions 3 and 4 to be parallel (to each other, not the wall) and of equal pressure.  This is the slip line condition.  Everything in regions 3 and 4 is parallel to the slip line.  The shocks at the wall reflect downstream of 3 & 4 because they can't continue parallel to the slip line -- they must turn to be parallel to the wall again.


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DFW International airport air traffic control
While listening to DFW air traffic control I hear pilots referring to Inwood and g bush when approaching the airport. What are these?

Related: https://aviation.stackexchange.com/questions/23743/what-is-the-difference-between-a-fix-a-waypoint-and-an-intersection/

Related: https://aviation.stackexchange.com/questions/17574/how-are-fixes-used-during-an-approach/

Related: https://aviation.stackexchange.com/questions/16171/what-is-a-navigational-fix-where-can-i-get-navigational-fixes-for-a-local-airpo

INWOD and GBUSH are RNAV fix points used for navigation. These two specific points are located on the ILS approach to runway 17L, on the extended centreline at 8.5 nautical miles (NM) and 5.4NM from the threshold respectively.
The names represent certain points in space defined by GNSS coordinates. Such 5-letter named points exists worldwide. The reason they exist is to make it easier for air traffic control and pilots to communicate, as they can just say a 5-letter word instead of having to spell out the GNSS coordinate (which would be a format like N43°38'19.39" W116°14'28.86"). For example, ATC might instruct pilots to "fly direct to GBUSH". It also makes it easier for pilots to input the position into the flight management system, as, again, they can just type in the 5 letters rather than typing in the coordinates.
The points do not correspond to any physical feature on the ground. They are simply virtual points defined by GNSS coordinates.
Below the approach plate for ILS 17L at KDFW where you can see INWOD and GBUSH depicted.

Source


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Left Bundle Block still a disqualifier?
Will a Left Bundle Block (lbbb) over 40 still disqualify you from passing the medical?

Assuming you are in the US and talking about the FAA, a left branch bundle block requires lab work and a radionuclide study.


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Why is there no Betz law in reverse?
Betz's law says that only 16/27 of the power in a wind stream can be converted into electrical power by a wind turbine. However, it appears that 100% of the power supplied to a fan motor can be converted into the power in a wind stream. Why is that? Or perhaps I am mistaken?

This doesn't appear to be a question about aviation as defined in the [help], but it might make a good question on [physics.se]

This is an excellent aerospace engineering question.  Over in Physics, they've probably never heard of it.  Betz's law is a really interesting consequence of the conservation of mass and momentum for a non-ducted turbine.  First, there are three velocities at play, far upstream of the turbine (largest) at the turbine, and far downstream of the turbine (smallest).  As it turns out, the velocity at the turbine is the average of the upstream and downstream velocities.  The power extracted is related to the overall change in velocity, but also the mass flow through the turbine.

As you vary the load (thrust/drag force) on the turbine, you change the velocity at the turbine, but since mass flow is rhovA (constant at all three locations) you also change the amount of mass flow captured.  As often happens in engineering, there are two effects -- the overall delta V, but also the size of the captured streamtube.  When these come together, they form an equation with a distinct maximum.

For a propeller, we can add as much power as we want -- and we are not normalizing it by the power contained in the streamtube ahead of the propeller as a reference.

I've voted to reopen since this is a question about aerodynamics as also @RobMcDonald highlighted

Betz's law is a really interesting consequence of the conservation of mass and momentum for a non-ducted turbine. First, there are three velocities at play, far upstream of the turbine (largest) at the turbine, and far downstream of the turbine (smallest). As it turns out, the velocity at the turbine is the average of the upstream and downstream velocities. The power extracted is related to the overall change in velocity, but also the mass flow through the turbine.
The Wikipedia article has a nice derivation of this, if it doesn't make sense to you, I can try to walk through it.
As you vary the load (thrust/drag force) on the turbine, you change the velocity at the turbine, but since mass flow is rhovA (constant at all three locations) you also change the amount of mass flow captured. As often happens in engineering, there are two effects -- the overall delta V, but also the size of the captured streamtube. When these come together, they form an equation with a distinct maximum.
For a propeller, we can add as much power as we want -- and we are not normalizing it by the power contained in the streamtube ahead of the propeller as a reference.


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How do I place wings on a fuselage for testing?
so I am researching the effects of wing geometry and anhedral angle on RDWs in ground effect using CFD. So far the research articles/papers I've read only test the wing without the fuselage as seen in the first image, link to the paper: https://doi.org/10.1016/j.cja.2022.06.015.

Now, if I were to use and test those wings with a fuselage how should I place it? Should I use method A or B (according to the bottom image, considering the circle/box to be the fuselage)? If I use method B, what should be the reference wing area?

Any advice is accepted, if you have sources to prove the right method that would be useful.

You're doing research. If your research is novel, then it's likely that nobody will know which is the "right" approach. Consider this in the context of a PhD defence and someone asks you why you chose the method you did, would "Somebody on the internet said so" be a satisfactory answer?

@CatchAsCatchCan I agree, I was hoping someone here would know and maybe have a source, or at least help lead me in the right direction since my major is not in aeronautical engineering, thanks for responding though!

@CatchAsCatchCan Do you have any recommended forums/books/journals where I could get advice/information regarding this, I am still an undergraduate and still getting into this so anything would help. Thanks!

The reference area is the one in A, i.e. considering the wing as extending inside the fuselage.

@CatchAsCatchCan any place is good to ask, even for doctoral studies. University cafeteria, the internet, train station toilet.

@sophit Ok thanks, I think It is generally agreed that the correct reference area is the one in A, but what about the anhedral? I think that B makes more sense structurally.

@RDD20: you can just choose one suitable reference and stick with it for all of your experiments.

Bottom line, the fuselage will reduce the cross-sectional area of the wing. My time in areo was long before CFD came along, but the principals will be the same. Without a fuselage, you are modeling lift across the entire wing. With the fuselage, the area taken up by the fuselage will no longer produce lift. Regardless of whether it is high/low wing (modify if fuselage is a "lifting-body" design). I would look for any research that discusses what fraction of the cross-sectional area of the fuselage to reduce lift/ground-effects by.

The conventional way to measure wing area includes the projected vertical area of the fuselage segment between them, either by spanwise line segments or by extending the leading and trailing edges (Peter Kämpf; Leeham News; FAI aeromodelling CIAM rules 2021, B.4.22 "When wings ... are built into the fuselage ... include that area contained within the normal contours of the [wing] extended so as to meet at the plane of symmetry.").
Thus, method B increases the wing area, which would be an unfair comparison to the wing alone.
I wouldn't call method A "right," but it does agree with the literature.

So I assume that the anhedral angle doesn't have to start at the centerline as long as the planform area remains the same? Although it would result in some differences with the experimental results from the paper.
Thanks for the answer btw!

That's right, to avoid comparing apples to oranges.  BTW, how stackexchange works: give us a few more days to post more answers, then click the checkmark to accept the one that best answered your question.

Method A and Method B are obviously going to give different results. Method C could be when the wings join below the fuselage and method D when the wings join above the fuselage.
   O
  / \        /O\
 /   \      /   \
Method C   Method D

In all four methods, there is only one where the fuselage does not have a significant impact on your results. If you want to research the impact of the fuselage, then you can not use that is the method. If you do not care about the impact of the fuselage, then it is the only method you can use.


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Why do cabin alt warning and takeoff config warning sound the same?
Having 2 same-sounding warnings is counterintuitive. Why did the cockpit designer do this?

In a modern airliner warning sounds are typically played by a computer via speakers in the cockpit. Adding additional sounds is easy, just add a file to the computer and done.
The Boeing 737 however was designed in the 1960s. Playing sounds via a computer simply wasn't an option. Of course there were technologies available in the 60s to record and re-play sounds (like vinyl records), but none of these were suitable for use in an airliner's cockpit. So any warning sound has to be physically produced instead of just played over a speaker.
For example, the fire alarm on the 737 uses the fire warning bell. The sound is created by the actual bell in the cockpit (although I'm not sure if the sound comes from speakers in modern variants of the 737). The horn was chosen for three different situations:

Condition: An intermittent or steady warning horn sounds:

In flight an intermittent horn indicates the cabin altitude is at or above 10 000 ft;
On the ground an intermittent horn indicates an improper takeoff configuration
when advancing thrust levers to takeoff thrust; and
In flight a steady horn indicates an improper landing configuration.

(Helios Airways Flight 522 - Final Report)
While it's not ideal to use the same horn for three different situations, these should never really occur simultaneously and the specific situation should therefore be recognizable from context. It was therefore deemed acceptable to re-use the same horn.
Otherwise, two additional distinct sounds would need to be created in the cockpit. This would require additional equipment to produce the sound, which increases cost, weight and (most importantly) complexity. Any additional component installed in the cockpit can fail, which might result in costly groundings until a repair can be made.
The final report of the Helios accident clearly lists this as a finding:

The use of the same aural warning to signify two different situations (Takeoff
Configuration and Cabin Altitude) was not consistent with good Human
Factors principles.

(Helios Airways Flight 522 - Final Report)
Boeing therefore installed two lights on the forward panel to distinguish between the two situations with intermittent horn:

Takeoff Configuration and Cabin Altitude Warning Lights

(Boeing 737 FCOMv2 - 15.10.3 Warning Systems)

Related: Aircraft Alerting Systems Standardization Study. Volume II. Aircraft Alerting System Design Guidelines (1981)


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Boeing 767 Engines Weight
The engines for the 767 weight about four tons each. That seems like a lot of weight for a lightweight aluminum airframe to carry in addition to passengers and other cargo. How is that possible, for a lightweight aluminum airframe to carry that much weight?

Metals (and fiber materials even more) are quite strong. Just think for example about a bike frame: it weights just a couple of kilos but can carry a hundred kg plus road roughness without any problem. Airframes are no exception to this rule.

Forget about the engines.  The 767-400 can carry up to 100 tons of fuel, most of which is actually stored in the wings.  The weight of the engine is small in comparison.

Forget the engines' weight. They are also producing over 20 tons of thrust each.

Aluminum is stronger than you might think, especially when it is configured in such a way to be more rigid.

How is that possible, for a lightweight aluminum airframe to carry that much weight?

Considering how much aluminum can handle weight-wise, 8 tons isn't that much for a 95 ton aircraft. Also, just because an airframe is light weight doesn't mean it's weak or can't handle a lot of stress. If you look at a bare 767 frame, you'll see how many aluminum ribs and reinforcements there actually are. (I was trying to find a good picture, but I didn't see any)
Also, the engines aren't really being held by the actual airframe. They're being held by mainly the wing spars and the wing box. To put it simply, the wing box and wing spars are separate from the aluminum airframe. The wing box and wing spars are actually supporting the whole aluminum airframe in flight.

Wing belongs to the airframe by definition, the engine(s) is attached to the airframe.

@sophit well yeah it is in some way attached to the airframe, through a set of linkages though.

@sophit  I think that what Wyatt meant to say was that the wing supports the whole fuselage in flight.

@MikeB yes that’s correct


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How to develop motion model for RC plane
I'm interested in controlling an RC plane autonomously, and to do that I need a basic motion model.
How would I do this? Particularly without a wind tunnel?

Constructing a mathematical dynamic system model of your aircraft is always the first task for a control engineer. In my oppinion (without claim to completeness) there are three ways to do this:

Analytically: For this, you normally start out by acquiring air-tables via either software (such as XFLR5, or a bit more basic XFOIL, or a CFD Software package) or wind-tunel data. The forces and moments acting on the aircraft can then be inserted into the well-known equations of motion of an aircraft, to acquire the dynamic system of an aircraft. This method is useful especially when the aircraft only exists on paper and has not been built (yet). However it requires some skill/money to acquire the air-data correctly.

Experimentally: Instead of relying on careful analysis, you can also just observe what the aircraft does in flight and do a system identification. For retrofitting an autopilot to an existing aircraft which already flies, this is significantly cheaper, easier and can be more exact then to analytically identifying the aircrafts's dynamics. The approach is as follows: You take your aircraft and fly a series of system identification maneuvers and record the flight data. This is easily done, because you need a data collection system anyways, as you want to outfit the aircraft with an autopilot. You then process this data via some math (I recommend this or this books, which should be some light reading (attention sarcasm)) to obtain your dynamic model. If done correctly, this method is extremly powerful and not only works for aircraft but also for rotorcraft. The US Army for example often uses CIFER, a commercial software toolbox based on MATLAB/Simulink. Because identified systems are built from flight data, it is guaranteed to be correct, which is always a big plus! Unfortunately, I am not aware of any publicly available software packages which do the heavy lifting of implementing system identification algorithms for you. However a lot of papers on this topic, especially for small UAVs exist, for example here or here. Especially the latter paper already includes transfer functions, perhaps these transfer functions already fit your airframe sufficiently well? If not, I think it is feasible to copy these methods with not too much work (although, I think some work will be necessary)...
Update Edit: Apparently the widely available Ardupilot software already contains a system identification mode which generates the necessary inputs and logs the aircraft response. As far as I can see you still have to implement the actual system identification algorithm yourself, however this at least makes the data collection easy.

Experimentally tuning: In this method, you take an easy to understand control structure (Most of the times this is a cascaded PID-controller) and tune the gains by hand until you are satisfied with the results. Normally you would need some intuition or do some analysis if your control structure is suitable for stabilizing the system in question, but for drones you can just use the typical cascaded PID controller.
This methods on the one hand is the most popular and on the other hand requires the least knowledge about your aircraft. For drones this method is dirt cheap and simple. Because of this 99% of the drones (and even some manned aircraft) are tuned this way. There are plenty of guides of how to do this on the internet, I just picked one at random here. Note however that empirical tuning does not give you any analytical stability margins, also the performance can be bad, depending on how good or bad you are at tuning your aircraft.

I love your idea of first developing a model for your aircraft before developing a controller, as it opens up the field of control engineering! There exist so many cool controllers, and everyone is just doing PID control, but constructing a good model can be quite a lot of work... Also, I am sorry for not giving you more concrete pointers of how to apply or develop methods 1 and 2, but then I would be sitting here tomorrow... I hope you find a good solution for your needs.

Yeah, allowing for MPC, trajectory optimization etc. I see PID loops everywhere, maybe it's because motion models are too complex or something idk

Btw how do you collect flight data for short range behavior? IMU gives some information but doesn't include absolute speed and can't be used for more than a couple seconds.

@FourierFlux You can use either a GPS sensor or pilot tubes for determining speed. Also AoA and Sideslip vanes can be used in addition to determine the speed vector. Often you just use all of these sensors for development. Please accept the answer, if you are satisfied...

@FourierFlux Perhaps a relatively interesting paper for you is the following one: https://ieeexplore.ieee.org/stamp/stamp.jsp?arnumber=7942783, or this one https://www.tandfonline.com/doi/full/10.1080/23311916.2022.2114196. You will find more similar examples on google.scholar.de by searching for UAV, fixed wing and system identification as keywords.


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Do electric hot air balloons exist?
With many things going electric these days, I wonder, is there an electric hot air balloon? One that doesn't use propane tanks to heat the balloon air.

I can't imagine that working very well. Batteries are heavy, and unlike propane (or similar fuels) the batteries don't get lighter as they discharge.

It really does seem a terrible idea.  I'm all for electric aircraft, but I think the best way to use batteries for a hot air balloon would be to set them on fire...

Electricity is relatively bad for heating things (compared e.g. to engines). OTOH you can use electricity on ground to generate fuels (PtX: Power to Gas/Liquid), which may be good because you can use the excess electricity on some time of most days (so also cheap)

It's not completely terrible. Ideas: helium-free tethered balloon, where power lines are easier than gas lines, or where payload power is beamed.

To a limited extent. There are electric fans for inflating hot air ballons, and electric maneuvering engines for balloons.
I haven't heard of anyone heating the air inside a balloon by electric means, although it seems easy enough to use an electric heat cannon for initial balloon inflation while on the ground. But even such cannons often combine propane heat with electric fans.
Using electricity to heat air isn't an environmental priority. Most electricity is produced by burning coal globally (although natural gas and nuclear have overtaken it in the most-developed countries). This heat-electricity-storage-heat conversion has an efficiency of 30%-50%, while burning propane directly, about 90%. The heater would also have to stay on the ground, supplied by AC grid power.
However, there does exist at least one solar-powered balloon. It doesn't use photovoltaics, but rather converts solar energy to heat directly, by being black on one side.

Perhaps one reason for limited innovation in hot air ballons is that a more environmentally sustainable alternative already exists: helium balloons. While helium is also in limited supply, coming as a byproduct of natural gas extraction, most of that helium isn't currently being collected.

Nope. the energy density of battery technology is way, way less than that of propane.  A battery source for heating the air inside a balloon would be far too heavy for the balloon to lift.

So much nope that even a (Chinese) patent for such a thing was allowed to quietly expire instead of paying the fee to preserve it.

I don't challenge your claims, but without figures to support them they look like opinions.

@CamilleGoudeseune Not saying the Chinese patent is or isn't practical, but many solid patents get abandoned for no better reason then they aren't commercially valuable enough to make paying fees reasonable.

First off, hot air balloons fly because we heat up the air inside that massive balloon (we call it the envelope), making the air less dense than the cooler air outside. This difference in density is what gets us floating up into the sky. Traditionally, we use propane burners—basically big ol' flamethrowers—to heat the air.
Now, could we swap out those propane burners for electric heaters? In theory, sure. But here's the thing: heating all that air requires a ton of energy. Electric heaters would need massive batteries, which are heavy. And adding heavy batteries to a balloon? Not exactly the best idea when you're trying to get airborne!
Developing an electric hot air balloon would involve a huge investment in research and technology to create lightweight, high-capacity batteries or super-efficient heaters. But here's the kicker: the hot air ballooning community is pretty small. We're a tight-knit group of enthusiasts, but there just aren't enough of us to make it a profitable venture for companies to pour resources into developing electric balloons.
So, while the idea is super cool, it's not really practical with our current technology and market size. Until there's a breakthrough that makes it feasible, we'll keep firing up the propane and enjoying those epic sunrise flights!


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For the same amount of thrust, is a jet airplane quieter than a propeller airplane?
If you have two airplanes, one generating thrust directly with a jet engine, and the other generating thrust with a propeller (from any of its power sources i.e: ICE/piston, electric, etc), which one is quieter?

It's hard to compare because the bulk of the noise from a turboprop will be around 60-100 Hz and will be dominant, while the jet engine will tend to be high intensity white noise (or partial white noise).  I think people perceive turboprops as being louder than jets.

The propeller driven aeroplane is quieter: it accelerates a higher mass at a lower velocity.
Aircraft thrust  is given as $ = \dot{m} \cdot (V_e - V_c)$, with

$\dot{m}$ = accelerated air mass flow in kg/s;
$V_e$ = average exhaust gas velocity;
$V_c$ = cruise speed.

Picture below is from this answer, which shows the difference in noise between a low bypass and a high bypass jet engine. The difference is caused by the engine on the right accelerating a higher flow to a lower velocity in order to obtain the same thrust, by means of its larger bypass fan. And a propeller is nothing more than an even larger bypass fan...

I agree with the logic but I’m not certain that a comparison between two engines with enclosed fans can be extended to an open-propeller engine.

Unfortunately, this answer misses the elephant in the room: With a propeller, you have a low count of blades at low rotation speed, where each blade's low/high pressure zone sends out a single crest/trough of a very strong sound wave. In principle, a turbine generates sound waves in the same way, but at much, much higher frequencies, easily going into the totally inaudible ultrasound range.

Another point is that the free blade ends of a propeller will generate strong vortices, which a shrouded impeller avoids entirely. So, reducing this question to the amount of air moved is a misleading oversimplification.

I don't believe this factors in noise from the piston engine either, does it?

There is noise, then again, there is vibration


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Does information about private heliports need to be published in the AIP?
I am seeking information on what international regulations specify regarding the publication of aeronautical data for a private heliports.

Only aerodromes (and heliports) participating to the international coordination of civil aviation, via the worldwide aeronautical fixed telecommunication network (AFTN), have to be listed in the AIP.
These coordinated aerodromes (and heliports) are open to any public operator, for civil traffic. They offer VFR and/or IFR services. Some are military.

The aeronautical information publication is a set of professional documents to facilitate international civil aviation under the recommendation of the ICAO, the international civil aviation organization. The documentation is described in annex 15 to the Chicago Convention. It includes aerodrome (and heliport) descriptions in section AD. Not all these aerodromes offer border controls, but they can be used for operations within the country after border controls have be satisfied. All airports in section AD are part of the AFTN network, and have an AFTN location identifier ("ICAO code").
However, the AIP is not concerned by private aerodromes/airfields/heliports where international public operators cannot land and take off. Note private is used in the sense of "restricted to private operations", as many privately owned aerodromes and airports are open to international aviation.
Aerodromes listed in the AIP cannot prohibit a public operator to use their facilities, except for technical limitations (e.g. the runway length or the material used to build it), or for commercial reasons (applicable fees not paid by the operator).
From the US AIP:

International arrivals with scheduled passenger service are not permitted to land at any aerodrome not listed in this AIP except in cases of real emergency or where special permission has been granted.

This mention is a copy of the one found in ICAO Doc 8126 which details the content of the aeronautical information services (AIS).

This doesn't preclude to maintain other directories. E.g. the FAA maintains a directory which goes beyond AIP ICAO requirements, and includes all airfields and navigation facilities declared to the FAA, for various reasons.

Unfortunately the FAA database for those airports is wildly out of date. When I attempted to update the status of some of the non-fields near me, I was told that they do not have the manpower to validate the entries. They also didn't seem to have a timeline on acting on the reports I made. I looked into it because a drone pilot friend was having his flight operations restricted by a private field that had been converted to a housing subdivision.


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What can cause an unstable phugoid mode?
For clarity, I am wondering about how aircraft design can lead to the phugoid poles becoming complex with a positive real part, so an oscillation with increasing amplitudes.
My understanding of the phugoid is limited to the reduced order approximation, with the damping ratio being inversely proportional to the lift-to-drag ratio of the aeroplane. As such, I have difficulty understanding what could cause a negative damping ratio. Is there any derivatives in particular that can cause an unstable phugoid? Where is the energy for the increasing amplitudes coming from?

The pilot itself can cause it or obviously some damage at the tailplane

@sophit With how long the period time of the phugoid typically is, PIO would have to be intentional, no? But, I was mostly wondering about ways aircraft design could create a naturally unstable phugoid. In what ways could a damaged tailplane cause an unstable phugoid?

Ah ok, you mean in the design phase. It's a very long mathematical story but the frequency of oscillation is mainly proportional to the inverse of the flying speed while the damping is mainly proportional to D/L. That implies that very efficient airplane (high L/D) flying at low speed (landing and taking off) might be problematic by a phugoid point of view. If this is what you're looking for I might elaborate it in a proper answer.

@sophit Yes, something akin to that is what I am looking for. I was aware of the approximations for the natural frequency and the damping ratio, but I don't understand how those can result in an unstable phugoid. Going by those, a high performance glider flying at low speeds should just result in a stable phugoid with very low damping and very high periods/very low frequencies, no? Does the approximation fail at those extremes or am I misunderstanding? I feel like I'm missing something and would be grateful for an answer.

Low speed implies high frequency (short period) and low damping, ideal conditions for PIO

@sophit Oh, you're right. I mixed those up. Low speeds means the phugoid mode has a high frequency. So it comes back to PIO? The phugoid doesn't become unstable without control inputs?

It might actually become unstable due to the thrust contribution but how exactly I don't remember  If you are not afraid of bug books with a lot of equation then "flight mechanics" is the keyword

You have to differentiate between stick fixed and stick free no?  A divergent phugoid stick fixed would be pretty bad. Stick free, I used to fly a glider (LS-1) that had an all flying tail with no antiservo tab to provide restorative trim forces. All of the AOA seeking trim forces came from the trim bungee spring in the control circuit. Letting it fly hands off, it was divergent with maybe a 15-30 sec period. I never let it go long enough to see if it would stall or dive past Vne.  Hands on the stick, it was fine.

@JohnK and others -- I routinely experience "stick-free" divergent pitch phugoids in several different gliders; I put "stick-free" in quotes because my hands are off the stick but friction is clearly not letting the stick move completely freely.  Watching it, it tends to move some near the max and min airspeed points, and then to freeze again.  I have let the process continue to the point where there was a marked stall break at the low-speed end of the oscillation and quite a high airspeed at the high-speed end of the oscillation, sometimes approaching the maneuvering speed, and felt that...

... and I felt that allowing the process to continue for a few more oscillations might indeed approach or exceed Vne.  A fun exercise is to get the oscillation going and then keeping hands off stick, try to use roll inputs (via the rudder pedals) to tame the oscillation-- it can be done, the trick is to try to have the bank angle shallowing (which tends to raise the nose above the horizon) at the same time that the airspeed is tending to increase, and try to have the bank angle increasing (which tends to lower the nose) at the same time that the airspeed is tending to decrease.

Gliders where I've seen these characteristics include Glassflugel 301 Libelle, and the lowly Schweizer 1-26.  As for true stick-fixed characteristics, I've found these hard to explore due to the difficulty of avoiding inadvertently avoiding making small pitch inputs.  Sometimes I've tried to do my best to accomplish that, and as far as I recall never had the impression that any initial pitch oscillation would not eventually damp out as the glider hunted for the "correct" aspd for the stick position-- at least so long as the first initial cycle didn't include a sharp stall break.


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Engine Failure - Immediately After Takeoff
Cessna 172 scenario: engine quits within seconds of takeoff at 200 ft AGL, necessitating a straight ahead crash landing. Assuming flaps 0, what is the optimal air speed to pitch for? Does it make any sense to use flaps?

Are you a student? What does your instructor say?

We try not to call it a "crash" landing, the aim here is just a normal landing without power!
As with all questions like this check your specific aircraft's POH, as that has the details for your model of C172.
Having got that obligatory line out the way, the best glide speed on my kneeboard for the 172 I fly in is 65 KIAS in the clean configuration. I know this because I brief it, along withwhat I'll do on an EFATO, every time I do my captains brief - and that includes to myself if I'm the only person on board! It's a good habit to get into, hopefully ensuring that if it does ever happen I'll just act instinctively.
Looking at a C172 POH Online it states exactly the steps you should take on page 3-5.

ENGINE FAILURE IMMEDIATELY AFTER TAKE OFF

Airspeed -- 65 KIAS (Flaps UP)
60 KIAS (Flaps DOWN)
Mixture -- IDLE CUT OFF
Fuel Selector Value -- OFF
Ignition Switch -- OFF
Wing Flaps -- AS REQUIRED
Master Switch -- OFF

So that answers your second question of whether it makes sense to use flaps - the answer is "if required".
From 200' you're going to be back on the ground in less than 30 seconds, and at a distance of something like 600 meters assuming zero wind. With a headwind that distance will be shorter. The most important thing you can do is, as always, aviate. Keep flying the aeroplane and do what is necessary to save your life.

Item (1) can be interpreted in different ways: (a) pitch for specific air speed depending on current flaps configuration, or (b) adjust the flaps as/if your air speed decreases to 60 KIAS. These are not mutually exclusive. What is the right way to interpret item (1)?
Second, how do I make sens of AS REQUIRED in item (5)?

@user443854 IMO, 1) is saying "if you are taking off clean, your best speed is 65, however if you're taking off with flaps it is 60". As for 5) if you need the extra drag to land back safely use flaps, otherwise don't - as per any landing really! In any case, it's all rather academic, as per my final paragraph, you have seconds, if not less, to think about this - just aviate. Worry about the debrief for what you should have done for the clubhouse/pub/pearly gates as appropriate. The key here is practice with an instructor and make it a memory action. You don't have time to think!

I’ve only flown 172s a couple of times and have never had an engine failure, so I’ll make a few observations from basic principles.  If you’re a wise pilot and are lucky enough to have enough runway left to stop then your first priority will be to get the nose down and maintain airspeed.  Then you can flare in to minimise sink rate as you touchdown.  If you have little space then you might keep your airspeed closer to stall, landing more heavily but with lower groundspeed.  If you’ve no clear ground to land on then it will be a snap decision as to what you want to hit and at what angle.  If you’ve just taken off it’s unlikely that you’ll have your flaps fully retracted.  I’m sure that others will have plenty more to contribute.

The optimal airspeed to pitch for, Vbg, around 65 knots flaps up, is the greatest glide distance.  200 feet/1 to 9 glide ratio is around 1/3 mile. (without headwind).
So, you're not going too far, even at Vbg.  It might be best to immediately prepare to land, as slow as possible.
If a suitable landing site is closer, use the short field landing procedure, go flaps 30, and approach at 55 knots.
Flaps lower the stall speed to 40 knots.  48 knots has 44 % more impact energy.
The only reason not to lower flaps is if a suitable landing area is available with a longer roll out, or if it is windy. The simplest way would be to pitch down to 65 knots for a zero flaps landing.